Thrust vector control with clustered nozzles



Dec. 20, 1966 R. SHORT ET AL 3,292,865

THRUS T VECTOR CONTROL WITH CLUSTERED NOZZLES Filed Oct. 17, 1963 INVENTOR United States Patent 3,292,865 THRUST VECTOR CONTROL WITH CLUSTERED NOZZLES Frederick R. Short and James P. Kern, Indianapolis, Ind.,

assignors to General Motors Corporation, Detroit, Mich, a corporation of Delaware Filed Oct. 17, 1963, Ser. No. 316,850

4 Claims. (Cl. 239-26525) This invention relates to a cluster of two or more convergent-divergent propulsion nozzles including a feature which permits directional control of the total thrust vectors.

The nozzle cluster feature comprises arranging the nozzles such that their divergent sections intersect, thereby eliminating portions of the divergent wall at which structural flow boundaries are replaced by gas dynamic boundaries. These gas dynamic boundaries of the expanding jets (which separate the adjoining jet streams) are automatically shifted by differential energization of the nozzles in a favorable direction to result in a jet stream angular deflection which produces a moment about the vehicle center of mass, which moment is additive to the moment resulting from the initiating differential axial thrust vectors. These controllable force moments thus become available for vehicle attitude and directional control. This feature, due to its high efiiciency, simplicity, and lack of moving nozzle parts has distinct advantages over other systems currently being used for thrust vector control.

Therefore, it is the object of the present invention to provide a cluster of two or more convergent-divergent propulsion nozzles which, through their particular configuration, permit directional control of the total thrust vector of the vehicle :being propelled.

Other objects, features, and advantages of the subject invention will become obvious to those skilled in the art to which the invention pertains upon study of the following detailed description of the preferred embodiment thereof and the attached drawings, wherein:

FIGURE 1 is a sectional view taken in the direction of arrows 1--1 in FIGURE 2;

FIGURE 2 is an end view taken in the direction of arrows 22 in FIGURE 1;

FIGURE 3 is a sectional view taken in the direction of arrows 33 in FIGURE 2; and

FIGURE 4 is a perspective View of the subject nozzle cluster configuration.

It should be clear that although but one embodiment of the subject invention is shown and described, many changes and modifications may be made thereto without departing from the scope of the invention.

FIGURE 2. illustrates one embodiment of this concept. Four convergent-divergent nozzles 10, 12, 14 and 16 oriented in a square pattern such that their divergent sections intersect at the plane where a flat barrier 18 interconnects the four nozzles. FIGURE 1 is an enlarged sectional view through the nozzles 10 and 12. The nozzles 10 and 12 are of the convergent-divergent type having a convergent inlet portion 20, a throat or venturi portion 22, and a diverging exit portion 24. An example of the nozzle cluster configuration shown may be four 16 to 1 area ratio convergent-divergent nozzles with their center intersection surface 18 located at approximately the six to one area ratio axial position.

In operation each of the nozzles 10, 12, 14 and 16 receives combustion gases from a gas source 25. Plugs 26 are provided to control the gas flow rate through each nozzle thereby resulting in varying nozzle inlet pressure levels. It is to be noted that this varying nozzle inlet pressure can be'obtained by using one combustion chamber (such as 25) and having a separate gas flow control- 3,292,865 Patented Dec. 20, 1966 ling plug (such as 26) for each nozzle, or by having four separate combustion chambers, each one feeding a specific exhaust nozzle. As seen in FIGURE 4, each of the four nozzles (10 thru 16) are integral with and converge cut of the combustion chamber housing 30. Thus, looking at FIGURE 1 again nozzle 10 has a remote wall 32 and an adjacent wall 34 which are formed integral with the combustion chamber housing 30 at its aft end. Likewise nozzle 12 has a remote wall 36 and an adjacent wall 38 also formed integral with the combustion chamber housing at its aft end. The term adjacent wall refers to that portion of the nozzle which lies between the intersections of the nozzle with the adjacent nozzles and which is the portion closest to the axis of the cluster. Likewise, the term remote wall refers to that portion of the wall which lies between the intersection points and which is farthest away from the vehicle axis. Nozzles 14 and 16 are also comprised of walls identical to those of 10 and 12 and also are formed integral with the combustion chamber housing 30 as is shown in FIGURE 3.

The remaining details of the structure of the device are not felt to be pertinent to the invention and will not be discussed. The operational details of the device will however be discussed in some detail. As mentioned previously, the inlet pressure to the nozzle cluster will be varied and as this pressure is varied the pressure level at the nozzle wall throughout the subsonic and supersonic expansion section of the particular nozzle will vary in proportion to the supply pressure change. With properly selected geometry, this wall pressure level variation will be approximately linear with the supply pressure level variation due to the nature of supersonic flow; however, this is not a necessary requirement.

In the subject embodiment, if the identical flow rate (inlet pressure) is supplied to each of the four nozzles, a resulting pure axial thrust action will be obtained with the pressure level at a given axial position being identical for all four nozzles. The internal flow boundaries between the four jet streams will be in the axial plane, thus there will be no resulting deflection of the vehicle.

Now if the supply pressure is increased in nozzles 14 and 16 and decreased in nozzles 10 and 12, the pressure levels in these nozzles will vary accordingly and the following thrust vectors will result. The combined axial thrust vectors of nozzles 14 and 16 will increase and their resultant will have its line of action to the outside of the nozzle center line. correspondingly, the axial thrust vectors of nozzles 10 and 12 will decrease and its resultant will also have its line of action to outside of the center line of these nozzles. Thus, the differential axial thrust between the resultant thrust of nozzles 14 and 16 and resultant vector of nozzles 10 and 12 will create 'a turning moment about the missile. Looking at FIGURE 4 this turning moment will cause the nozzle assembly to rotate to the right thus going into the plane of the drawings away from the viewer. In addition, a side thrust is created due to the differential wall pressure ofnozzles 14 and 16 as compared to the pressures of nozzles 10 and 12. This differential pressure level acts over the unbalanced area bounded by surfaces 46, 44 and 18 in FIGURE 1. Since the pressure is higher in nozzles 14 and 16 the resulting side force is in the direction to the right when looking at the FIGURE 2 and creates a turning moment about the rocket which will be additive to the turning moment caused by the axial thrust vector moment. An added axial thrust vector is produced by the pressure acting on the surface 18 and is additive to the axial thrust vectors of nozzles 14 and 16 and of nozzles 10 and 12.

In further explanation of the resulting internal flow pattern, the flow boundary between the jet stream is deflected towards nozzles 10 and 12. Thus, this increases theeffective expansion ratio for nozzles 14 and 16 .and

decreases the efiective expansion area ratio for nozzles and 12. For definite values of ambient pressure, these changes arebothin the desirable direction since the a turning moment in the other direction. The invention also lends itself to many other changes which will increase the adaptability of the device. In the embodiment shown the center lines of the nozzles are allshown parallel to the missile axis. Further vector force potential could be gained by canting or tilting the nozzle center line in radial planes. Thus, in such a configuration the center lines of the nozzles would tend to converge toward the missile axis.

Controllable roll vectors could be obtained in several ways, such as by tilting the center lines of the nozzles in non-radial planes, or by circumferentially stationing nozzles in other than equal angular increments. The first of these, the tilting of the center lines of the nozzles in nonradial planes, would be accomplished by tilting the axis of the nozzles in planes parallel to the missile axis. This would cause the axes of the nozzles to tend to diverge or converge towards each other. The second method of controlling the roll vectors mentioned, circumferentially spacing the nozzles in other than equal angular increments, would result in adjacent nozzles being either more than or less than ninety degrees apart.

It should be clear then that desirable clusters of more than or less than four nozzles can :be made using the subject invention, and that vectoring can be obtained in any desired direction by proper variations in the supply pressures of the various intersecting nozzles. The overall expansion area ratios of the nozzles and the area ratios at which they intersect'can be varied over wide ranges to produce desired optional configurations for specific applications. In passing it should be mentioned that a greater percentage of the missile frontal area can be utilized at nozzle exit area when this intersecting concept is used in comparison with an ordinary nozzle cluster, and this would be a definite advantage particularly in an upper stage of a multi-stage missile which would be operating during the upper altitude portions of the rocket flight.

Although the subject invention has been described in but one specific embodiment it should be clear to those skilled in the art to which the invention pertains that some of the changes mentioned and many other changes and modifications may be made to the subject device without departing from the scope of the invention.

We claim: V v 1. An exhaust duct configuration for a rocket engine comprising a cluster of at least three convergent-divergent exhaust nozzles, each of said exhaust nozzles having its own converging portion and its own throat portion,

each of said exhaust nozzles intersecting the adjacent exhaust nozzles in the diverging portion of the nozzles,

each of said exhaust nozzles having a portion of its adjacent wall in the diverging portion of the nozzle removed such that each of said nozzles" diverges into a common duct formed by the remote walls of each of i said exhaust nozzles.

2. A thrust level and thrust vector controlling system for a rocket engine, comprising a cluster of scarfed exhaust nozzles each having an individual and separate throat portion, said scarfing of said cluster, of exhaust nozzles being the result of the absence of a portion :of the adjacent wall of each of said exhaust nozzlessuch that each nozzle intersects the adjacent nozzles and the adjacent walls thus terminate, and each said nozzle of said cluster of nozzles diverges into a commonduct formed by the remote walls of each of said nozzles, and individually controllable means for supplying hot exhaust gas to each nozzle of said cluster of exhaust nozzles,

the variation of flow of hot gases to any given exhaust nozzle resulting in a different resultant thrust vector f01' the engine. 7

3. A thrustlevel and thrust vector controlling system for a rocket engine, comprising a cluster of scarfed exhaust nozzles each having an individual and separate throat portion, said scarfing of said cluster of exhaust nozzles being the result of the absence of a portion of the adjacent wall of each of said exhaust nozzles such that each nozzle intersects the adjacent nozzles and the adjacent walls thus terminate, and each said nozzle of said cluster of nozzles diverges into a common duct 1 of said exhaust noz-. zles, means for supplying hot exhaust gas to each noz-' formed by the remote walls of each zle of said cluster of exhaust nozzles, and individually controllable plug means for controlling the supply of, hot gases to each nozzle of said cluster of exhaust noz-,

zles, the variation of flow of hot gases to any given exhaust nozzle resulting in a different resultant-thrust vector for the engine.

4. An exhaust duct configuration as recited in claim 1 in which said adjacent wall of each nozzle upstream of said removed portion terminates in a plane perpendicular to the thrust axis of the nozzle, cluster, and in:

'cluding a barrier extending between said adjacent wall portions closing the interspace between the said nozzles.

References Cited by the Examiner UNITED STATES PATENTS MARK NEWMAN, Primary Examiner. A. L. SMITH, Assistant Examiner. 

1. AN EXHAUST DUCT CONFIGURATION FOR A ROCKET ENGINE COMPRISING A CLUSTER OF AT LEAST THREE CONVERGENT-DIVERGENT EXHAUST NOZZLES, EACH OF SAID EXHAUST NOZZLE HAVING ITS OWN CONVERGING PORTION AND ITS OWN THROAT PORTION, EACH OF SAID EXHAUST NOZZLES INTERSECTING THE ADJACENT EXHAUST NOZZLES IN THE DIVERGING PORTION OF THE NOZZLES, EACH OF SAID EXHAUST NOZZLES HAVING A PORTION OF ITS ADJACENT WALL IN THE DIVERGING PORTION OF THE NOZZLE REMOVED SUCH THAT EACH OF SAID NOZZLES DIVERGES INTO A COMMON DUCT FORMED BY THE REMOTE WALLS OF EACH OF SAID EXHAUST NOZZLES. 